Compressor Rotor for Gas Turbine Engine With Deep Blade Groove

ABSTRACT

A compressor rotor has a rotor centered on an axis, and a groove with opposed side edges. The groove receives a plurality of removable compressor blades, and has tangential sides. The tangential side surfaces are in contact with said tangential sides of the groove, and at least one slot is cut into the side edges. A first radial distance is measured from a radially outer edge of the side edge to a radially outer beginning point of the tangential sides of the groove. A second radial distance is radially between the radially outer beginning point of the tangential sides to a radially inner end of the tangential sides of the groove. A ratio of the first and second radial distance is between 1.1 and 5.0.

BACKGROUND

This application relates to a rotor for use in a compressor for a gas turbine engine, wherein a blade groove has tangential sides that are at a relatively deep location.

Gas turbine engines are known, and typically include a fan delivering air into a compressor. The air is compressed and delivered into a combustor where it is mixed with fuel and ignited. Products of this combustion pass downstream over the turbine rotors, driving them to rotate.

The compressors typically include a plurality of rotors, some of which include grooves that receive removable blades. The blades extend from a root radially outwardly to an airfoil. The grooves are provided with lock and load slots which are utilized to load the blades into the grooves. Typically there are tangential sides to the grooves, which provide a reaction surface for the sides of the blades. These tangential sides transmit mechanical tensile stress back into the rotor. These tensile stresses offset thermally induced compressive stresses, which are particularly concentrated at such slots as load or lock slots.

SUMMARY

In a featured embodiment, a compressor rotor has a rotor centered on an axis, and a groove with opposed side edges. The groove receives a plurality of removable compressor blades. The groove has tangential sides. The blades have tangential side surfaces to be in contact with the tangential sides of the groove. At least one slot is cut into the side edges. A first radial distance is defined measured from a radially outer edge of the side edge to a radially outer beginning point of the tangential sides of the groove. A second radial distance is defined radially between the radially outer beginning point of the tangential sides to a radially inner end of the tangential sides of the groove. A ratio of the first radial distance to the second radial distance is between 1.1 and 5.0.

In another embodiment according to the previous embodiment, the tangential sides of the groove are defined at an angle. The angle is between 0 and 75 degrees.

In another embodiment according to any of the previous embodiments, the slots include lock slots and load slots to assist in loading blades into the grooves.

In another embodiment according to any of the previous embodiments, a bearing surface slot is formed within at least one of the tangential sides.

These and other features of this application will be best understood from the following specification and drawings, the following of which is a brief description.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 schematically shows a gas turbine engine.

FIG. 2A shows a first feature of a compressor rotor.

FIG. 2B shows another feature.

FIG. 3A is a cross sectional view through a rotor.

FIG. 3B is a view similar to FIG. 3A, but with a blade removed.

DETAILED DESCRIPTION

A FIG. 1 schematically illustrates a gas turbine engine 20. The gas turbine engine 20 is disclosed herein as a two-spool turbofan that generally incorporates a fan section 22, a compressor section 24, a combustor section 26 and a turbine section 28. Alternative engines might include an augmentor section (not shown) among other systems or features. The fan section 22 drives air along a bypass flowpath B in a bypass duct defined within a nacelle 15, while the compressor section 24 drives air along a core flowpath C for compression and communication into the combustor section 26 then expansion through the turbine section 28. Although depicted as a turbofan gas turbine engine in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to use with turbofans as the teachings may be applied to other types of turbine engines including three-spool architectures.

The engine 20 generally includes a low speed spool 30 and a high speed spool 32 mounted for rotation about an engine central longitudinal axis A relative to an engine static structure 36 via several bearing systems 38. It should be understood that various bearing systems 38 at various locations may alternatively or additionally be provided.

The low speed spool 30 generally includes an inner shaft 40 that interconnects a fan 42, a low pressure compressor 44 and a low pressure turbine 46. The inner shaft 40 may be connected to the fan 42 through a geared architecture 48 to drive the fan 42 at a lower speed than the low speed spool 30. The high speed spool 32 includes an outer shaft 50 that interconnects a high pressure compressor 52 and high pressure turbine 54. A combustor 56 is arranged between the high pressure compressor 52 and the high pressure turbine 54. A mid-turbine frame 57 of the engine static structure 36 is arranged generally between the high pressure turbine 54 and the low pressure turbine 46. The mid-turbine frame 57 further supports bearing systems 38 in the turbine section 28. The inner shaft 40 and the outer shaft 50 are concentric and rotate via bearing systems 38 about the engine central longitudinal axis A which is collinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 then the high pressure compressor 52, mixed and burned with fuel in the combustor 56, then expanded over the high pressure turbine 54 and low pressure turbine 46. The mid-turbine frame 57 includes airfoils 59 which are in the core airflow path. The turbines 46, 54 rotationally drive the respective low speed spool 30 and high speed spool 32 in response to the expansion.

The engine 20 in one example is a high-bypass geared aircraft engine. In a further example, the engine 20 bypass ratio is greater than about six (6), with an example embodiment being greater than ten (10), the geared architecture 48 is an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3 and the low pressure turbine 46 has a pressure ratio that is greater than about 5. In one disclosed embodiment, the engine 20 bypass ratio is greater than about ten (10:1), the fan diameter is significantly larger than that of the low pressure compressor 44, and the low pressure turbine 46 has a pressure ratio that is greater than about 5:1. Low pressure turbine 46 pressure ratio is pressure measured prior to inlet of low pressure turbine 46 as related to the pressure at the outlet of the low pressure turbine 46 prior to an exhaust nozzle. The geared architecture 48 may be an epicycle gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.5:1. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present invention is applicable to other gas turbine engines including direct drive turbofans.

A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. The fan section 22 of the engine 20 is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft, with the engine at its best fuel consumption—also known as “bucket cruise Thrust Specific Fuel Consumption (‘TSFC’)”—is the industry standard parameter of lbm of fuel being burned divided by lbf of thrust the engine produces at that minimum point. “Low fan pressure ratio” is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about 1.45. “Low corrected fan tip speed” is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram ° R)/(518.7 ° R)]^(0.5). The “Low corrected fan tip speed” as disclosed herein according to one non-limiting embodiment is less than about 1150 ft/second.

FIG. 2A shows a portion of a rotor 80 which may be incorporated into a compressor section in a gas turbine engine such as gas turbine engine 20. As shown, there are tangential sides 82 radially within a groove 200 defined between rim edges 88.

A bearing surface slot 86 is shown within the side 82.

FIG. 2B shows opposed rim edges 88, a load slot 92, and lock slots 90, all of which are cut into the facing edges 88.

When blades are mounted within the groove 200, and the rotor is driven at high speed, edges of the blades contact the tangential sides 82, and transmit mechanical tensile stresses. As mentioned above, those mechanical tensile stresses can offset thermally induced compressive stresses, which are concentrated in the slots.

FIG. 3A is a cross-sectional view through the rotor 80. As shown, sides 83 of a blade 94 contact the sides 82. An airfoil 95, partially illustrated, extends radially outwardly. As the rotor 80 is driven to rotate at high rates of speed, the blade is urged radially outwardly due to centrifugal forces, and the mechanical tensile stresses from sides 83 contacting the tangential sides 82 become high.

The present invention addresses this concern by making the sides 82 relatively radially deep compared to the prior art.

As shown in FIG. 3B, a first distance d₁ can be defined between a radially outer point 89 of the side edge 88, and extending inwardly to a radially inner end 91 of the side edge 88. These distances are measured relative to a center axis A. The tangential sides 82 begin at point 91 and extend radially inwardly to point 93. In one embodiment the tangential side extends at an angle A. The radial distance between point 91 and 93 is d₂. This is a radial distance and not the length along the surface of side 82.

In the prior art, a ratio of d₁ to d₂ was much smaller than it is in the present rotor 80. In the prior art a ratio of d₁ to d₂ may have been approximately 0.9.

In this application a ratio of d₁ to d₂ was between 1.1 and 5.0. The angle A was 45 degrees in one embodiment. In embodiments, A may be between 0 and 75 degrees.

With the increased depth of the tangential sides 82, the thermally induced compressive stresses are offset by greater mechanical tensile stresses, and are therefore not as concentrated in slots as in the prior art.

Although an embodiment of this invention has been disclosed, a worker of ordinary skill in this art would recognize that certain modifications would come within the scope of this invention. For that reason, the following claims should be studied to determine the true scope and content of this invention. 

1. A compressor rotor comprising: a rotor centered on an axis, and having a groove with opposed side edges, and said groove receiving a plurality of removable compressor blades, said groove having tangential sides, and said blades having tangential side surfaces to be in contact with said tangential sides of said groove, and at least one slot cut into said side edges; and a first radial distance defined measured from a radially outer edge of said side edge to a radially outer beginning point of said tangential sides of said groove, and a second radial distance defined radially between said radially outer beginning point of said tangential sides to a radially inner end of said tangential sides of said groove, and a ratio of said first radial distance to said second radial distance being between 1.1 and 5.0.
 2. The compressor rotor as set forth in claim 1, wherein said tangential sides of said groove are defined at an angle, said angle being between 0 and 75 degrees.
 3. The compressor rotor as set forth in claim 1, wherein said slots include lock slots and load slots to assist in loading blades into the grooves.
 4. The compressor rotor as set forth in claim 1, wherein a bearing surface slot is formed within at least one of said tangential sides.
 5. The compressor rotor as set forth in claim 2, wherein said slots include lock slots and load slots to assist in loading blades into the grooves.
 6. The compressor rotor as set forth in claim 5, wherein a bearing surface slot is formed within at least one of said tangential sides.
 7. The compressor rotor as set forth in claim 2, wherein a bearing surface slot is formed within at least one of said tangential sides.
 8. The compressor rotor as set forth in claim 3, wherein a bearing surface slot is formed within at least one of said tangential sides.
 9. A gas turbine engine including: a fan delivering air into a bypass duct and into a compressor, a combustor, and a turbine section; and the compressor including a compressor rotor having a rotor centered on an axis, and having a groove with opposed side edges, and said groove receiving a plurality of removable compressor blades, said groove having tangential sides, and said blades having tangential side surfaces to be in contact with said tangential sides of said groove, and at least one slot cut into said side edges, a first radial distance defined measured from a radially outer edge of said side edge to a radially outer beginning point of said tangential sides of said groove, and a second radial distance defined radially between said radially outer beginning point of said tangential sides to a radially inner end of said tangential sides of said groove, and a ratio of said first radial distance to said second radial distance being between 1.1 and 5.0.
 10. The gas turbine engine as set forth in claim 9, wherein said tangential sides of said groove are defined at an angle, said angle being between 0 and 75 degrees.
 11. The gas turbine engine as set forth in claim 9, wherein said slots include lock slots and load slots to assist in loading blades into the grooves.
 12. The gas turbine engine as set forth in claim 9, wherein a bearing surface slot is formed within at least one of said tangential sides.
 13. The gas turbine engine rotor as set forth in claim 10, wherein said slots include lock slots and load slots to assist in loading blades into the grooves.
 14. The gas turbine engine as set forth in claim 13, wherein a bearing surface slot is formed within at least one of said tangential sides.
 15. The gas turbine engine as set forth in claim 10, wherein a bearing surface slot is formed within at least one of said tangential sides.
 16. The gas turbine engine as set forth in claim 11, wherein a bearing surface slot is formed within at least one of said tangential sides. 